|Publication number||US8166767 B2|
|Application number||US 12/905,663|
|Publication date||May 1, 2012|
|Filing date||Oct 15, 2010|
|Priority date||Sep 15, 2006|
|Also published as||CA2601934A1, CA2601934C, US7836702, US20100236257, US20110023499|
|Publication number||12905663, 905663, US 8166767 B2, US 8166767B2, US-B2-8166767, US8166767 B2, US8166767B2|
|Inventors||Nicolas Grivas, Bhawan B. Patel, Oleg Morenko, Eric Durocher|
|Original Assignee||Pratt & Whitney Canada Corp.|
|Export Citation||BiBTeX, EndNote, RefMan|
|Patent Citations (31), Non-Patent Citations (2), Referenced by (6), Classifications (19), Legal Events (2)|
|External Links: USPTO, USPTO Assignment, Espacenet|
The present application is a divisional application of U.S. patent application Ser. No. 11/532,231 filed Sep. 15, 2006 now U.S. Pat. No. 7,836,702, the entire content of which is incorporated herein by reference.
The present invention relates generally to gas turbine engines, and more particularly to the interface between a high pressure turbine vane platform and an exit duct of a combustor therein.
The secure and well sealed inter-engagement of the combustor and the high pressure turbine (HPT) stage immediately downstream therefrom is important in a gas turbine engine. This portion of the engine experiences the highest temperatures, and therefore the joint formed between the combustor and the HPT vane assembly must be able to accommodate thermal deflections of the two structures due to the high temperatures. Thermal growth mismatch between the combustor walls and the platforms of the HPT vane may occur for a variety of reasons, such as a result of differences in materials, structure, temperature variations, surface temperatures and gas flow patterns for example. Accordingly, the joints linking these two structures must be able to accommodate these deflections, without reducing sealing effectiveness. A tight seal therebetween during engine operation is required to prevent high pressure air, typically fed to the regions around the combustor, from leaking into the lower pressure gas flowing through the turbine section.
However, achieving an ideal clearance between the small exit duct (SED) and the HPT vane platform can be difficult, as too tight a clearance can cause binding between the two parts such that relative movement therebetween is limited and therefore thermal growth mismatch cannot be readily absorbed by a suitable joint arrangement, while too large a clearance gap therebetween can result in hot gas leakage and a reduction in the pressure differential required for impingement airflow cooling of the hot SED surfaces.
It is accordingly an object of the present invention to provide an improved interface between a high pressure turbine vane platform and a combustor flange in a gas turbine engine.
Therefore, in accordance with the present invention, there is provided an assembly for a gas turbine engine, the assembly comprising a combustor and a vane assembly downstream from the combustor, the vane assembly including a plurality of airfoils radially extending between annular inner and outer platforms defining an annular gas path therebetween, at least the outer platform forming a first sliding joint connection with an adjacent outer combustor wall such that relative axial displacement therebetween is permitted, a radial clearance gap being defined between the outer platform and the outer combustor wall, said radial clearance gap being minimized at engine operating temperature such that radial sealing between the outer platform and the outer combustor wall is substantially provided, and wherein the outer platform includes a plurality of depressions located in an outer circumferential surface thereof opposite said outer combustor wall, said depressions being disposed in regions of expected higher thermal growth about the circumference of the outer platform such that thermal growth of the outer platform is substantially uniform circumferentially therearound and thus said minimized radial clearance gap is maintained during engine operation to prevent binding contact between the outer platform and the outer combustor wall.
There is also provided, in accordance with the present invention, an assembly for a gas turbine engine, the assembly comprising: a combustor having a combustor wall circumscribing a gas path therewithin and an exit duct end, the exit duct end having an outer flange portion radially spaced from the combustor wall in a direction away from the gas path to form an annular slot; a vane assembly disposed downstream of the combustor, the vane assembly including a plurality of airfoils extending between annular inner and outer platforms; and a portion of at least the outer platform extending axially within said annular slot and defining a sliding joint connection between the combustor and the vane assembly, the sliding joint connection permitting relative axial displacement between the vane assembly and the combustor, said portion of the outer platform including a plurality of depressions located in an outer circumferential surface thereof opposite said outer flange portion, said depressions being disposed in regions of expected higher thermal growth about the circumference of the outer platform such that thermal growth of the outer platform is substantially uniform circumferentially therearound.
There is also provided, in accordance with the present invention, a combustor-to-vane joint assembly for a gas turbine engine, the combustor including an annular exit end for mating with an annular vane ring assembly, the vane ring assembly including a plurality of airfoils radially extending between inner and outer platform rings, the joint comprising: inner and outer female retaining members integral with and defining a downstream end of the annular exit, the female retaining members each including spaced-apart annular portions; and inner and outer male insertion members integral with and defining an upstream end of the respective inner and outer platform rings of the vane ring assembly; the vane assembly being mounted to the combustor solely by insertion of the inner and outer male insertion members between the spaced-apart portions of the respective inner and outer female retaining members; and wherein at least the outer male insertion member of the outer platform ring has a plurality of depressions located in an outer circumferential surface thereof opposite at least one of said spaced-apart annular portions of said outer female retaining member, said depressions being disposed in regions of expected higher thermal growth about the circumference of the outer platform ring such that thermal growth of the outer platform ring is substantially uniform circumferentially therearound.
There is also provided, in accordance with the present invention, a method of providing an interface between a combustor and a vane assembly in a gas turbine engine, the method comprising: providing a vane assembly mounted to an exit duct of the combustor, the vane assembly having a plurality of airfoils radially extending between annular inner and outer platforms, at least the outer platform cooperating with a flange portion of the combustor exit duct to form an axial sliding joint therebetween; forming a plurality of depressions in a portion of the at least the outer platform opposite the flange portion, said depressions being disposed in local regions of expected higher thermal growth about the circumference of the at least the outer platform such that thermal growth of the entire outer platform is substantially uniform circumferentially therearound; determining a maximum amount of thermal growth differential exhibited between the flange portion and the at least the outer platform over an engine operating temperature range; and defining a radial clearance gap between the flange portion and the at least the outer platform at ambient temperature that is slightly greater than the determined maximum amount of thermal growth differential over the engine operating temperature range, such that the radial clearance gap is minimized but maintained during the engine operating temperature range thereby providing substantial radial sealing between the at least the outer platform and the combustor while preventing binding contact therebetween.
Further features and advantages of the present invention will become apparent from the following detailed description, taken in combination with the appended drawings, in which:
The turbine section 18 further comprises at least a first, or high pressure (HP), turbine stage 17 which is immediately downstream from the combustor 16. The high pressure turbine (HPT) stage 17 includes a turbine rotor (not shown) with a plurality of radially extending turbine blades, and a HPT vane assembly 22 (see
An integral upstream lug portion 44 of the vane inner platform 26 extends upstream from the leading edge 21 of the vane airfoil 24, and projects into an annular slot defined between an outer flange portion 31 and an inner flange portion 29 of the combustor inner wall 30. The inner joint assembly 34 accordingly formed, permits the upstream lug portion 44 of the vane inner platform 26 to be relatively axially displaceable within the flange portions 29, 31 of the combustor inner wall 30. Particularly, the inner joint assembly 34 is a sliding-type joint, which permits the upstream lug portion 44 of the vane inner platform 26 to axially slide within the flange portions of the combustor inner wall 30 in a direction substantially parallel to the direction 25 of the hot gas flow (i.e. parallel to the central longitudinal axis 11 of the engine 10). The inner joint assembly 34 may also permit a small radial displacement of the upstream lug portion 44 relative to the inner and outer flange portions 29 and 31 of the combustor inner wall 30 in the form of a radial thermal growth differential generated therebetween.
Referring now to the outer joint assembly 36 between the HPT vane assembly 22 and the upstream combustor 16, an upstream end or lug portion 42 of the vane outer platform 28 projects upstream from the leading edge 21 of the vane airfoil 24, for engagement with the outer wall 32 of the combustor 16. The combustor's outer wall 32 comprises an inner flange portion 33, which is preferably integrally formed with the main wall of the combustor, and an outer flange portion (or finger portion) 35 which together define annular space-apart portions of the outer combustor wall 32 at exit of the combustor 16. The inner flange portion 33 and the outer flange portion 35 define a radial gap or slot 39 therebetween. The upstream lug portion 42 of the vane outer platform 28 freely extends within the radial slot 39, such that relatively displacement between the upstream lug portion 42 and the combustor outer wall 32 is possible. Particularly, the upstream lug portion 42 is able to axially slide therewithin. Accordingly, the outer joint assembly 36 is also a sliding-type joint, and provides a sealing joint which is able to accommodate thermal deflection of the parts without requiring any locating fasteners such as rivets or bolts.
During operation of the gas turbine engine 10, the hot combustion gases flowing through the annular gas path passage 23 cause thermal growth of the HPT vane assembly 22, including the outer and inner platforms 28, 26 thereof, particularly in the outer radial direction 50. While one would typically expect corresponding thermal growth of the combustor walls, at least the outer flange portion 35 of the combustor outer wall 32 generally does not expand as much as the outer platform 28 of the vane assembly, as it is not in direct contact with the hot combustion gases and is exposed to the cooler high pressure air fed to the region 40 surrounding the combustor 16. Accordingly, the outer flange portion 35 experiences lower thermal growth in a radially outward direction 50 than does the upstream end 42 of the outer platform 28. This resulting thermal growth differential causes a reduced radial clearance gap 52 in the sliding outer joint assembly 36 during engine operation, particularly between the outer flange portion 25 of the combustor outer wall 32 and the upstream end 42 of the vane outer platform 28.
However, too much thermal growth of the platforms of the HPT vane assembly relative to the combustor walls can cause binding to occur in the sliding joints therebetween. Conversely, too little can leave a radial clearance gap 52 therebetween which is large enough to allow leakage therethrough. Unless a reasonably efficient radial seal is created between the outer and inner platforms 28, 26 and the corresponding combustor walls 32, 30, the high pressure air fed to the region 40 around the combustor 16 can leak into the lower pressure gas path 23 flowing through the HPT stage 17 of the turbine section 18. The upstream end 42 of the vane outer platform 28 has a radially outer surface 43 which opposes an inwardly facing surface 54 of the aligned portion the outer flange portion 35 of the combustor outer wall 32. The radial clearance gap between the outer vane platform 28 and the outer combustor wall 32 is therefore radially defined between the outer surface 43 and the inwardly facing surface 54 thereof.
Therefore, when the gas turbine engine is in operation, the radial clearance gap 52 defined between the outer platform 28 and the opposed outer wall of the combustor is ideally minimized (i.e. kept as small as possible) such as to provide a substantial radial seal between the two components. However, in order to prevent any binding therebetween, which can occur if this radial clearance gap 52 is eliminated completely (for example, if thermal growth of the outer platform is significantly greater than the clearance gap 52 in a cold or ambient temperature), at least a small clearance gap 52 is also preferably maintained through the range of engine operating temperatures.
However, this need to select a cold radial clearance gap 52 sufficient to allow for a minimized but maintained clearance gap at hot engine operating temperatures is made further difficult by the fact that different regions of the vane platforms experience different amounts of thermal growth. For example, either due to higher local temperature pockets, different geometries, different cooling airflow patterns, etc., certain regions of the vane platforms 28, 26 experience greater thermal growth than others, particularly in the radial direction 50.
Referring now to
The radial gap 52 between the outer platform 28 and outer combustor wall 32 (i.e. the so called “small exit duct” portion of the combustor 16) can therefore be optimized, such that it is minimized but not eliminated completely during the range of engine operating temperatures. Accordingly, the “cold” size of the radial gap 52 (i.e. at ambient temperature) is selected such that, upon an expected amount of radial thermal growth when exposed to the predetermined engine operating temperatures, the “hot” size of the radial gap 52 will be small enough to limit significant gas leakage therethrough while remaining sufficient large enough to prevent binding between the outer platform and the opposed combustor flange 35.
Therefore, in order to provide such an interface between the combustor and the vane assembly in the gas turbine engine, a plurality of the depressions 62 are first formed in at least the outer platform 28, opposite the combust exit flange portion 35 of the outer combustor wall 32 and within the regions 60 of expected higher thermal growth. Accordingly, the thermal growth of the entire outer platform 28 will be substantially uniform circumferentially therearound when exposed to the hot engine operating temperatures. Next, a maximum amount of thermal growth differential exhibited between the flange portion 35 and the outer platform 28 over an entire range of the engine operating temperatures is determined. Then, the cold radial clearance gap 52 between the flange portion 35 and the outer platform 28 at ambient temperature is defined, such that it is slightly greater than the determined maximum amount of thermal growth differential over the engine operating temperature range. This accordingly permits the radial clearance gap 52 at hot engine operating temperatures to be minimized, but nonetheless maintained, thereby providing substantial radial sealing between the vane assembly and the combustor while preventing binding contact therebetween.
The depressions 62, in at least one embodiment, include flats in the otherwise curved annular outer surface 54, which may be formed by any suitable method such as machining, casting or the like. Although the depressions 62 are likely only very slight (for example, flat regions formed into the curved outer circumferential surface of the platform's upstream end), if more extreme differences are expected in the thermal growth of these regions relative to the circumferentially spaced about regions between each airfoil, then the depressions 62 within regions 60 may be made more dramatic (for example having a greater, and therefore perhaps more visually noticeable, depth within the outer platform).
Although the radial clearance gap 52 between the outer vane platform 28 and the outer combustor wall 32 is described above, a similar radial clearance gap between the inner vane platform 26 and the inner combustor wall 30 similarly exists within the inner sliding joint 34. Similarly, the upstream end 44 of the inner platform 26 also may include similar depressions 62 therein in regions 60 of expected greater thermal growth, as described above relative to the outer platform.
The sliding joint connections between both the vane outer and inner platforms 28, 26 and the corresponding combustor walls 32,30 improve the engagement between the HPT vane assembly 22 and the combustor 16 at engine operating temperature. In comparison with typical rigidly fixed prior art combustor-HPT vane joints, the outer and inner joint assemblies 34,36, and particularly the outer joint assembly 36, comprise fewer parts, require no permanent fasteners such as rivets or bolts, and provide improved sealing as a result of the thermal growth differential between the vane outer platform 28 and the cooler combustor outer flange portion 35. In all gas turbine engines, and particularly those intended for aerospace applications, weight and cost considerations are of prime importance. Accordingly, reducing the number of parts and simplifying their design is a constant goal of all aerospace designers and engineers. The combustor sliding joint assemblies 34,36 achieve this by providing sealing connections which have fewer, less complex parts, thereby greatly simplifying the connection between the HPT vane assembly 22 and the combustor 16, particularly at the outer joint assembly 36 between the vane outer platform 28 and the combustor outer wall 32.
The embodiments of the invention described above are intended to be exemplary. Still other modifications are available, and those skilled in the art will therefore appreciate that the forgoing description is illustrative only, and that various alternatives and modifications can be devised without departing from the spirit of the present invention. Accordingly, the present invention is intended to embrace all such alternatives, modifications and variances which fall within the scope of the appended claims.
|Cited Patent||Filing date||Publication date||Applicant||Title|
|US3286461||Jul 22, 1965||Nov 22, 1966||Gen Motors Corp||Turbine starter and cooling|
|US3302926||Dec 6, 1965||Feb 7, 1967||Gen Electric||Segmented nozzle diaphragm for high temperature turbine|
|US3529906||Oct 30, 1968||Sep 22, 1970||Westinghouse Electric Corp||Static seal structure|
|US3670497||Sep 2, 1970||Jun 20, 1972||United Aircraft Corp||Combustion chamber support|
|US3965066||Mar 15, 1974||Jun 22, 1976||General Electric Company||Combustor-turbine nozzle interconnection|
|US4011718 *||Aug 1, 1975||Mar 15, 1977||United Technologies Corporation||Gas turbine construction|
|US4195476||Apr 27, 1978||Apr 1, 1980||General Motors Corporation||Combustor construction|
|US4635332||Sep 13, 1985||Jan 13, 1987||Solar Turbines Incorporated||Sealed telescopic joint and method of assembly|
|US4668164||Dec 21, 1984||May 26, 1987||United Technologies Corporation||Coolable stator assembly for a gas turbine engine|
|US4702670||Mar 20, 1987||Oct 27, 1987||Rolls-Royce||Gas turbine engines|
|US4798514||Jun 24, 1980||Jan 17, 1989||Rolls-Royce Limited||Nozzle guide vane structure for a gas turbine engine|
|US4821522||Jul 2, 1987||Apr 18, 1989||United Technologies Corporation||Sealing and cooling arrangement for combustor vane interface|
|US4889469||Jul 28, 1978||Dec 26, 1989||Rolls-Royce (1971) Limited||A nozzle guide vane structure for a gas turbine engine|
|US4920742||May 31, 1988||May 1, 1990||General Electric Company||Heat shield for gas turbine engine frame|
|US5078576||Jun 29, 1990||Jan 7, 1992||Rolls-Royce Plc||Mounting system for engine components having dissimilar coefficients of thermal expansion|
|US5398496||Mar 9, 1994||Mar 21, 1995||Rolls-Royce, Plc||Gas turbine engines|
|US5407319||Mar 9, 1994||Apr 18, 1995||Rolls-Royce Plc||Sealing structures for gas turbine engines|
|US5417545||Mar 3, 1994||May 23, 1995||Rolls-Royce Plc||Cooled turbine nozzle assembly and a method of calculating the diameters of cooling holes for use in such an assembly|
|US5470198||Aug 31, 1994||Nov 28, 1995||Rolls-Royce Plc||Sealing structures for gas turbine engines|
|US6269628||Jun 10, 1999||Aug 7, 2001||Pratt & Whitney Canada Corp.||Apparatus for reducing combustor exit duct cooling|
|US6860108||Jan 22, 2003||Mar 1, 2005||Mitsubishi Heavy Industries, Ltd.||Gas turbine tail tube seal and gas turbine using the same|
|US6895757||Feb 10, 2003||May 24, 2005||General Electric Company||Sealing assembly for the aft end of a ceramic matrix composite liner in a gas turbine engine combustor|
|US7000406 *||Dec 3, 2003||Feb 21, 2006||Pratt & Whitney Canada Corp.||Gas turbine combustor sliding joint|
|US7350358 *||Nov 16, 2004||Apr 1, 2008||Pratt & Whitney Canada Corp.||Exit duct of annular reverse flow combustor and method of making the same|
|US7546743 *||Oct 12, 2005||Jun 16, 2009||General Electric Company||Bolting configuration for joining ceramic combustor liner to metal mounting attachments|
|US20050120718||Dec 3, 2003||Jun 9, 2005||Lorin Markarian||Gas turbine combustor sliding joint|
|US20080061515 *||Sep 8, 2006||Mar 13, 2008||Eric Durocher||Rim seal for a gas turbine engine|
|US20090208326||Apr 20, 2009||Aug 20, 2009||Eric Durocher||Rim seal for a gas turbine engine|
|GB1089660A||Title not available|
|GB1091573A||Title not available|
|GB2102897A||Title not available|
|1||Pratt & Whitney Canada JT15D-5F Maintenance Manual, selected pages from Chapter 72-50-Turbine Section.|
|2||Pratt & Whitney Canada JT15D-5F Maintenance Manual, selected pages from Chapter 72-50—Turbine Section.|
|Citing Patent||Filing date||Publication date||Applicant||Title|
|US9512740||Nov 22, 2013||Dec 6, 2016||Siemens Energy, Inc.||Industrial gas turbine exhaust system with area ruled exhaust path|
|US9540956||Nov 22, 2013||Jan 10, 2017||Siemens Energy, Inc.||Industrial gas turbine exhaust system with modular struts and collars|
|US9587519||Nov 22, 2013||Mar 7, 2017||Siemens Energy, Inc.||Modular industrial gas turbine exhaust system|
|US9598981||Nov 22, 2013||Mar 21, 2017||Siemens Energy, Inc.||Industrial gas turbine exhaust system diffuser inlet lip|
|US9631517||Dec 29, 2012||Apr 25, 2017||United Technologies Corporation||Multi-piece fairing for monolithic turbine exhaust case|
|US9644497||Nov 22, 2013||May 9, 2017||Siemens Energy, Inc.||Industrial gas turbine exhaust system with splined profile tail cone|
|U.S. Classification||60/772, 415/138, 60/799, 60/796, 60/800|
|International Classification||F01D25/26, F02C7/20|
|Cooperative Classification||F05D2240/80, F05D2250/41, Y02T50/675, F01D11/005, F05D2260/94, F01D9/023, F23R2900/00005, F05D2260/941, F01D9/042|
|European Classification||F01D11/00D, F01D9/02B, F01D9/04C|
|Jan 11, 2012||AS||Assignment|
Owner name: PRATT & WHITNEY CANADA CORP., CANADA
Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:GRIVAS, NICOLAS;DUROCHER, ERIC;MORENKO, OLEG;AND OTHERS;SIGNING DATES FROM 20060907 TO 20060913;REEL/FRAME:027516/0291
|Oct 27, 2015||FPAY||Fee payment|
Year of fee payment: 4